With reference to FIG. 1, a ducted fan gas turbine engine generally indicated at 10 has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, and intermediate pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23.
The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 13 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.
The performance of gas turbine engines, whether measured in terms of efficiency or specific output, is improved by increasing the turbine gas temperature. For any engine cycle compression ratio or bypass ratio, increasing the turbine entry gas temperature produces more specific thrust (e.g. engine thrust per unit of air mass flow). It is therefore desirable to operate the turbines at the highest possible temperatures. However, as turbine entry temperatures increase, the life of a turbine generally shortens, necessitating the development of better materials and/or the introduction of improved cooling systems. One group of improved materials includes so-called ceramic matrix composite, CMC, materials. CMCs offer superior temperature and creep resistant properties for gas turbine engines and have a considerably lower density than their superalloy counterparts making them ideal for aeroengines. Further, because they have a higher temperature tolerance, CMC materials require less cooling which acts to increase specific fuel consumption further.
CMC materials generally consist of ceramic fibres embedded with a ceramic body. There are different materials available for fibres and body. Two of the more promising materials for gas turbine engines are silicon carbide fibres within a body of silicon carbide, so-called SiC/SiC, and aluminium oxide fibres within an aluminium oxide body, which is referred simply as an oxide CMC. The processes for manufacturing CMC materials are reasonably well known and understood in the art.
FIG. 2 shows a high pressure turbine section of the engine shown in FIG. 1. Thus there is shown an nozzle guide vane 212 and turbine blade 214 in flow series having aerofoil sections within the main gas path 216. The turbine blade includes a tip 218 which is radially shrouded by a seal segment 220. The seal segment 220 bounds and defines the main gas path 216 on the outboard side of the turbine core. The seal segment 220 in the example shown is manufactured from a CMC material so as to provide some of the advantages outlined above.
The seal segment 220 includes a radially inboard gas washed surface 222 with radially extending supporting walls 224 which project towards and append from the engine casing via an intermediate support structure in the form of a carrier structure 226. The walls 224 include forward facing hooks which mate with corresponding formations on the carrier 226. The carrier 226 is attached to the engine casing 229. FIG. 2 shows a single seal segment 220 in streamwise section but it will be appreciated that this is one of many circumferentially arranged seal segments 220 configured to provide an annulus around the turbine wheel.
The seal segments 220 are separated by an intersegment gap which allows for relative movement between the seal segments 220 when in use. The intersegment gap is provided by opposing circumferential end faces.
It is well known that such intersegment gaps require sealing to prevent a flow of cooling air from the outboard side into the main gas path 216, or the vice versa. Thus, it is known to include slots 230 in the intersegment gap faces to receive seal strips. The seal strips sit in the seal strip slots and restrict the flow passage between the main gas path and outboard side of the seal segment. The seal strips and slots 230 are of suitable dimensions to accommodate some relative movement between adjacent seal segments.
The present invention seeks to provide an improved sealing system for seal segments, and in particular, CMC seal segments. Additionally, the invention seeks to provide a seal arrangement which may be useful for sealing between other intercomponent gaps such as those between nozzle guide vane platforms for example.